The present invention relates to aircraft with electric flight controls comprising a fuselage able to deform and vibrate longitudinally and laterally with the formation of vibration nodes and antinodes distributed along the longitudinal axis of said aircraft. It relates quite particularly to long-length airplanes which have high longitudinal flexibility. However, it advantageously applies equally well to airplanes of a shorter length and lower flexibility.
It is known that an aircraft with electric flight controls has flight controls such as sticks, mini sticks, rudder bars, etc., which are equipped with electric transducers so that they generate electric flight control datums representative of the action that a pilot exerts on them. It also comprises a flight control computer which, on the basis of said electric flight control datums generated by said flight controls and of flight control parameters originating, for example, from sensors, formulates electric commands that said flight control computer applies to actuators tasked with moving the control surfaces of the aircraft.
It is also known that aircraft with electric flight controls are provided with an inertial reference system (generally known as an IRS) comprising elements useful in navigation, such as the inertial unit, and elements useful in flight control, such as gyrometers and accelerometers. Finally, it is known that all these elements, whether they have to do with navigation or flight control, are grouped together in an IRS unit arranged at a given point on said aircraft. Of course, as a result, this IRS unit is subjected to the action of the deformations of the fuselage, which deformations occur mainly along the axes of pitch and yaw under the effect of the turning of the control surfaces or the effect of external disturbances.
Because of the high time constant attached to the elements useful in navigation, such deformations have only a small action thereon. By contrast, in order to get around the problems of interaction between the deformations of the fuselage and the elements useful in flight control, it is essential to have filtering means on the control surface control lines.
However, in the case of aircrafts with high longitudinal flexibility, the deformations become greater, which means that it is then necessary to perform extremely intense filtering of said control lines, and this introduces significant phase shifts thereinto and therefore detracts greatly from the performance of said control lines.
It is an object of the present invention to overcome this drawback.
To this end, according to the invention, an aircraft with electric flight controls, provided with control surfaces able to be moved by electrically operated actuators, said aircraft comprising:
controls and at least one flight control computer, said controls being actuated by a pilot and generating electric flight control datums which are sent to said flight control computer, the latter computer generating, on the basis of said electric flight control datums and flight control parameters, commands in roll, pitch and yaw, which are sent to said actuators to move said control surfaces;
an inertial reference system comprising elements useful in navigation and elements useful in flight control, the latter elements being either of the gyrometer type or the accelerometer type; and
a fuselage able to deform and vibrate with the formation of vibration nodes and antinodes distributed along the longitudinal axis of said aircraft,
is notable in that:
said inertial reference system has an exploded structure with said elements useful in flight control separated from said elements useful in navigation;
said elements useful in flight control are distributed along said fuselage;
each element useful in flight control, of the gyrometer type, is arranged at a vibration node of said fuselage;
each element useful in flight control, of the accelerometer type, is arranged at a vibration antinode of said fuselage; and
said elements useful in flight control are connected to said flight control computer so that the measurement signals they deliver are used as flight control parameters.
Thus, said accelerometers allow the measurement of the accelerations of the aircraft including vibrational movements of the fuselage, while the gyrometers allow the measurement of the rotation rates without incorporating the structural modes of the fuselage thereinto. These accelerometer and gyrometer measurements are sent to the flight control computer which in consequence formulates commands for the control surfaces.
The flight control laws incorporated into this computer therefore do not need to filter the vibrational movements of the fuselage. This is because the structural modes measured by the accelerometers can be actively checked by the flight control laws while the gyrometers do not measure deformations of the fuselage. In the most frequent scenario, the aircraft fuselage deforms and vibrates in such a way as to have a vibration antinode at each of its ends, and a vibration node near its center of gravity.
In this case, said aircraft comprises:
at least one front accelerometer arranged at the front part of said fuselage and delivering a vertical acceleration measurement and a lateral acceleration measurement;
at least one rear accelerometer arranged at the rear part of said fuselage and delivering a vertical acceleration measurement and a lateral acceleration measurement; and
at least one gyrometer arranged near said center of gravity of said aircraft and delivering roll rate, pitch rate and yaw rate measurements.
It is then advantageous for:
said vertical acceleration measurements generated by said front accelerometer and by said rear accelerometer respectively and the pitch rate measurement generated by said gyrometer, to be used as flight control parameters to formulate said pitch commands;
said lateral acceleration measurements generated by said front accelerometer and by said rear accelerometer respectively, and said roll rate and yaw rate measurements generated by said gyrometer, to be used as flight control parameters to formulate said roll commands; and
said lateral acceleration measurements generated by said front accelerometer and by said rear accelerometer respectively, and said roll rate and yaw rate measurements generated by the gyrometer, to be used as flight control parameters for formulating said yaw commands.
The aircraft can then comprise:
means of filtering said acceleration measurements and the rate measurement or measurements to eliminate measurement noise therefrom and avoid spectrum folding;
gain multipliers for weighting each of said filtered acceleration or rate measurements;
phase control means for said filtered and weighted acceleration measurements; and
summing means for summing said filtered, weighted and phase-controlled acceleration measurements, said filtered and weighted rate measurement or measurements and the corresponding electric flight control datum to formulate the corresponding command.
The aircraft may also, for formulating roll and yaw commands, comprise means of integrating the roll rate so as to create information about the roll angle, which information is sent to said summing means after it has been weighted by a gain multiplier.
Of course, in such an architecture, all the gains are optimized so as to satisfy the compromises between performance and stability. It is also found that the architecture according to the present invention makes it possible to dispense with low-frequency filters, even though the aircraft might be very flexible.